Nearly all satellites utilize solar panel systems to harvest electricity for powering on-board systems, charging batteries, etc. The utilization of solar energy in satellites often requires a tradeoff between maximizing size of the solar panels and minimizing valuable launch vehicle real estate. For example, when a satellite is prepared to be launched into orbit via a spacecraft, the associated solar panel equipment may be secured in a retracted position to conserve space and to avoid damage to the equipment or the deployer, and to mitigate the chances of orbital debris caused by deployment. Once the satellite is in orbit, the solar panels are unfolded and deployed to an operative position.
FIG. 1 depicts a top-view example representation of a prior art satellite 102 with an “accordion”-style solar panel array 104 arrangement in stowed (A), unfolding (B), and deployed (C) positions. In such systems, the array 104 solar panels are pivotally coupled end-to-end, folded accordion style for stowage (A), and secured with tie-down/release mechanisms 106, 108 along both edges of the array and/or at a central portion of the panels. Once the satellite is in orbit, the restraining mechanisms are released to allow the array 104 to unfold into a substantially planar (or linear) configuration to receive photons from the sun.
In many prior art systems, the array 104 is attached at a midpoint 110 on the body of the satellite 102. In such systems, a half-width panel 112 may be utilized to allow for folding the array 104 against one side of the satellite 102 during stowage (A), but such an arrangement may leave a portion of unused volume 114 between the folded array 104 and the body of the satellite 102. Furthermore, the use of a reduced size panel 112 (due to the mid-point attachment 110) may reduce the effective size (and power harvesting potential) of the unfolded array 104.
One of the drawbacks with accordion-style arrangements and/or attachment points of the conventional system is that the edges of the stowed panel 104 typically need to be secured on both respective sides (and/or the center) of the satellite body by the tie-down/release mechanisms 106, 108, which can create extra bulk. In such prior art systems, the chances for a deployment fault can be increased due to the added coordination needed for releasing both tie-down/release mechanisms 106, 108. In some systems, one or more of the tie-down/release mechanisms 106, 108 may become detached during deployment, and may create additional orbital debris and/or may damage the panel 104 or satellite 102.
A need exists for improved systems and methods to address such challenges.